EP2820250B1 - Moteur à turbines à gaz équipé d'une section d'induction assemblée à la soufflante - Google Patents

Moteur à turbines à gaz équipé d'une section d'induction assemblée à la soufflante Download PDF

Info

Publication number
EP2820250B1
EP2820250B1 EP13784828.9A EP13784828A EP2820250B1 EP 2820250 B1 EP2820250 B1 EP 2820250B1 EP 13784828 A EP13784828 A EP 13784828A EP 2820250 B1 EP2820250 B1 EP 2820250B1
Authority
EP
European Patent Office
Prior art keywords
gas turbine
turbine engine
fan
engine according
inducer
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
EP13784828.9A
Other languages
German (de)
English (en)
Other versions
EP2820250A2 (fr
EP2820250A4 (fr
Inventor
Frederick M. Schwarz
Daniel Bernard KUPRATIS
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2820250A2 publication Critical patent/EP2820250A2/fr
Publication of EP2820250A4 publication Critical patent/EP2820250A4/fr
Application granted granted Critical
Publication of EP2820250B1 publication Critical patent/EP2820250B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D15/00—Adaptations of machines or engines for special use; Combinations of engines with devices driven thereby
    • F01D15/12—Combinations with mechanical gearing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/072—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with counter-rotating, e.g. fan rotors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/047—Heating to prevent icing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/40—Transmission of power
    • F05D2260/403—Transmission of power through the shape of the drive components
    • F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
    • F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type

Definitions

  • This disclosure relates to a gas turbine engine with a fan-tied inducer section.
  • a typical jet engine has multiple shafts or spools that transmit torque between turbine and compressor sections of the engine.
  • a low speed spool generally includes a low shaft that interconnects a fan, a low pressure compressor, and a low pressure turbine.
  • the low pressure turbine drives the low shaft, which drives the low pressure compressor.
  • a geared architecture connects the low shaft to the fan. Air exiting the fan at the root has relatively low energy, which generates a swirling effect that makes it difficult to efficiently feed air into the low pressure compressor.
  • a gas turbine engine having the features of the preamble of claim 1 is disclosed in US 2010/0126141 A1 or EP 2071153 A2 .
  • the present invention provides a gas turbine engine as set forth in claim 1.
  • the at least one inducer stage comprises one or more inducer blades fixed for rotation with the fan rotor and a core inlet stator fixed to a non-rotating engine structure.
  • the core inlet stator is positioned axially between the fan and the inducer blades.
  • the core stator is positioned aft of the blades.
  • the at least one inducer stage comprises a plurality of inducer stages coupled to the fan rotor.
  • each inducer state comprises one or more blades fixed for rotation with the fan rotor and a core stator fixed to a non-rotating engine structure.
  • the compressor is positioned immediately aft of the at least one inducer stage.
  • the at least one inducer stage is positioned forward of the speed reduction device.
  • the compressor comprises a low pressure compressor
  • the gas turbine engine includes a high pressure compressor positioned aft of the low pressure compressor and driven by a second shaft positioned radially outwardly relative to the shaft that drives the low pressure compressor.
  • the core inlet stator is a variable vane.
  • the core inlet stator is heated for anti-icing.
  • the one or more inducer blades are heated for anti-icing.
  • the fan rotor turns in the same direction as the shaft. In an alternative embodiment, the fan rotor turns in an opposite direction from the shaft.
  • the speed change device comprises a gearbox including a sun gear in meshing engagement with star gears and a ring gear in meshing engagement with the star gears, and wherein the ring gear at drives the fan.
  • the speed change device comprises a gearbox including a sun gear in meshing engagement with a plurality of planetary gears supported by a planet carrier, and a ring gear in meshing engagement with the planet gears, and wherein the ring gear is fixed and the planet carrier provides input to the fan.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26 then expansion through the turbine section 28.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26
  • the engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • the inner shaft 40 is connected to the fan 42 through a speed reduction device, such as a geared architecture 48 for example, to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A, which is collinear with their longitudinal axes.
  • the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
  • the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • the engine 20 in one example is a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10)
  • the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • the engine 20 bypass ratio is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5:1.
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 22 of the engine 20 is designed for a particular flight condition -- typically cruise at about 0.8 Mach and about 35,000 feet (10668 m).
  • 'TSFC' Thrust Specific Fuel Consumption
  • Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45.
  • the "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about 1150 ft / second (350.5 m/s).
  • a gas turbine engine 60 shown in Figure 2 includes a two-spool turbofan as described above, which generally incorporates a fan section 22, a compressor section 24 with high 52 and low 44 pressure compressors, and a turbine section 28 with high 54 and low 46 pressure turbines.
  • the low pressure turbine 46 is comprised of a plurality of stages.
  • the low pressure turbine 46 includes a first stage 62, a second stage 64, and a third stage 66.
  • the high pressure turbine 54 is comprised of a first stage 68 and a second stage 70 that are positioned forward of the plurality of stages 62, 64, 66 of the low pressure turbine 46.
  • Each of the stages for the high 54 and low 46 pressure turbines includes a plurality of blades coupled to a respective rotor.
  • blades 71 of the low pressure turbine 46 are coupled to a first rotor 72 and blades 74 of the high pressure turbine 54 are coupled to a second rotor 76.
  • the first rotor 72 is configured to drive the low shaft 40 and the second rotor 76 is configured to drive the high shaft 50.
  • Each stage of the high 54 and low 46 pressure turbines also includes a plurality of vanes (not shown) interspersed with the blades where the vanes are mounted to a static engine structure 36.
  • the high pressure compressor 52 is comprised of first 80, second 82, third 84, and fourth 86 stages.
  • the low pressure compressor 44 is comprised of first 88, second 90, third 92, and fourth 94 stages that are positioned forward of the plurality of stages 80, 82, 84, 86 of the high pressure compressor 52.
  • Each of the stages of the high pressure compressor 52 includes a plurality of blades 96 that are coupled to a rotor 98 that is driven by the high shaft 50.
  • Each of the stages of the low pressure compressor 44 is comprised of blades 100 that are coupled to a rotor 102 that is driven by the low shaft 40.
  • Each stage of the high 52 and low 44 pressure compressors also includes a plurality of vanes (not shown) interspersed with the blades where the vanes are mounted to a static engine structure 36.
  • Various bearings 38 rotatably support the high 50 and low 40 shafts as known.
  • the fan section 22 includes a fan 42 that is driven by the geared architecture 48.
  • the fan 42 is comprised of a plurality of fan blades 104 that are coupled to a fan rotor 106 for rotation about the axis.
  • the geared architecture 48 couples the low shaft 40 to the fan rotor 106 such that the fan rotor 106 rotates at a lower speed than the low shaft 40.
  • the geared architecture 48 is an epicyclic gear arrangement that includes a plurality of star or planet gears driven by a sun gear fixed for rotation with the low shaft 40. The star gears drive a ring gear that is configured to drive the fan rotor 106.
  • the engine 60 also includes an inducer section 110 that comprises a fan-tied compressor stage, i.e. the inducer section is an additional compressor stage that is connected to the fan rotor 106.
  • the inducer section 110 serves to efficiently feed the low pressure compressor 44 a more controlled/stabilized air flow.
  • the inducer may allow the manufacturer to make one engine model without the inducer and use the exact same core on an engine model with the inducer. This approach would be a way of improving manufacturing efficiencies since the engine core with the inducer could in turn be used at higher thrust while maintaining the same peak core temperatures due to the higher air flows provided by the additional inducer stage or stages. In this way most of the engine part numbers would be common between the two models, thereby eliminating duplication in engineering work, development work, toolings, and other savings.
  • the inducer section 110 includes at least one inducer stage 112 that is driven by the fan rotor 106.
  • the inducer stage 112 comprises one or more blade rows 114 fixed for rotation with the fan rotor 106 and a core stator structure 116 fixed to the non-rotating static engine structure 36.
  • the core stator structure 116 is configured to facilitate reducing swirl coming off of the fan and diffusing the air flow.
  • the core stator structure 116 includes one or more vanes 118 fixed to the static engine structure 36. Additional support for the stator structure 116 is provided by a connection to a strut 120.
  • a bearing 122 rotatably supports the shaft 40 for rotation relative to the strut 120 and stator structure 116.
  • a thrust bearing 148 also provides support for the fan and inducer assembly.
  • the core stator structure 116 is positioned axially between the fan 42 and the blades 114 of the inducer stage 112.
  • the core stator structure 116 is positioned aft of the blades 114 of the inducer stage 112.
  • the at least one inducer stage comprises a plurality of inducer stages 112, 130 coupled to the fan rotor 106.
  • Figure 4 shows an example configuration having a first inducer stage 112 and a second inducer stage 130 that is configured similarly to the first inducer stage 112.
  • the second inducer stage 130 comprises one or more blades 132 fixed for rotation with the fan rotor 106 and a core stator structure 116 fixed to the static engine structure 36 as described above.
  • the core stator structure 116 can optionally include a second set of vanes 134 positioned aft of the blades 118 of the first inducer stage 112 and forward of the blades 132 of the second inducer stage 130.
  • the vanes could be positioned respectively aft of each set of blades in a manner similar to that shown in Figure 3 .
  • the geared architecture 48 comprises a gearbox.
  • Figures 5A and 5B shows two different examples of gearboxes.
  • Figure 5A shows a gearbox that comprises a star gear configuration. This configuration includes a sun gear 150 in meshing engagement with star gears 152 and a ring gear 154 in meshing engagement with the star gears 152.
  • the sun gear 150 is driven by the shaft 40 and the ring gear 154 at drives the fan rotor 106.
  • Figure 5B shows a gearbox that comprises a planetary gear configuration.
  • This configuration includes a sun gear 160 in meshing engagement with a plurality of planetary gears 162 supported by a planet carrier 164 and a ring gear 166 in meshing engagement with the planet gears 162.
  • the ring gear 166 is fixed to a static structure 36 and the planet carrier 164 provides input to the fan rotor 106.
  • the gearbox defines a gearbox axial center-plane P (see Figure 2-4 ).
  • the low pressure compressor 44 is positioned immediately aft of the inducer section 110 and the inducer section 110 is positioned forward of the gearbox axial center-plane P.
  • the low pressure compressor 44 is positioned aft of the gearbox axial center-plane P.
  • the various configurations described above provide a geared turbofan with a slow turning, fan-tied auxiliary compressor stage or stages, and a separate higher speed low pressure compressor/low pressure turbine that are tied to a common single shaft.
  • the fan-tied low pressure compressor stage or stages provide an inducer that is connected to the fan rotor immediately aft of the fan. This provides several benefits.
  • the configurations disclosed above improve engine operability by fractionally reducing the pressure rise required of the higher speed low pressure compressor and moving that fractional pressure rise to the lower speed fan rotor utilizing the associated inducer stage.
  • Moving low pressure compressor stages from the higher speed low pressure compressor shaft to the slower rotating fan rotor improves engine operability by reducing the inertia of the higher speed low pressure compressor and turbine, which are tied to the same single shaft.
  • the inertia of the stages in an inducer configuration is decreased by a factor of 1/GR2 where GR2 is the square of the speed reduction ratio of the gear. For example, if the gear ratio GR is 2, locating a stage as fan-tied reduces the inertia of that low compressor stage by the factor of 0.25 relative to locating the stage within the higher speed low pressure compressor.
  • one or more of the core inlet stator vanes 118 could be a variable vane that rotates along a spanwise axis (as schematically indicated by dashed arrow 170 in Figure 2 ) in order to better align the airfoil to the input flow. This would be especially desirable in reducing takeoff peak temperatures in the core.
  • the core inlet stator vanes 118 and/or the inducer blades 114, 132 could be heated for anti-icing conditions.
  • a heating device is schematically shown at 174 in Figure 4 .
  • the fan 42 and inducer section 110 turn in the same direction as the fan drive turbine using a speed reduction device or a gearbox of the planet type as shown in Figure 5B .
  • the sun gear 160 provides the input torque
  • the ring gear 166 is fixed to the supporting structure 36 and the carrier 164 of the gears 162 between the sun 160 and the ring gear 166 is connected to the fan hub 106 and provides the rotational torque required by the fan 42.
  • the fan-tied low pressure compressor enables even more pressure to be addressed by the fan rotor 106 as the fan-tied low pressure compressor more easily accommodates more work being done by the fan rotor 106 than the counter rotating high speed low pressure compressor does without the presence of a fan-tied compressor stage. This also increases the supercharging temperature of the high speed low pressure compressor and, thus, results in a lower tip Mach number for the first rotor of the high speed low pressure compressor, resulting improved efficiency.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Claims (16)

  1. Moteur à turbine à gaz comprenant :
    un arbre (40) définissant un axe de rotation ;
    un dispositif de réduction de vitesse (48) entraîné par l'arbre (40) ; et
    une soufflante (42) incluant un rotor de soufflante (106) entraîné par le dispositif de réduction de vitesse (48) ;
    caractérisé en ce qu'il comprend en outre :
    un compresseur (44) comprenant une pluralité d'étages de compresseur entraînés par l'arbre ; et
    au moins un étage inducteur (112, 130) positionné à l'arrière de la soufflante (42) et en amont du compresseur (44) et accouplé pour rotation avec le rotor de soufflante (106).
  2. Moteur à turbine à gaz selon la revendication 1, dans lequel l'au moins un étage inducteur (112, 130) comprend une ou plusieurs pales d'inducteur (114, 132) fixées pour rotation avec le rotor de soufflante (106) et un stator d'entrée central (116) fixé sur la structure de moteur non rotative (36).
  3. Moteur à turbine à gaz selon la revendication 2, dans lequel le stator d'entrée central (116) est positionné de manière axiale entre la soufflante (42) et les pales d'inducteur (114).
  4. Moteur à turbine à gaz selon la revendication 2, dans lequel le stator central (116) est positionné à l'arrière des pales (114).
  5. Moteur à turbine à gaz selon la revendication 2, 3 ou 4, dans lequel le stator d'entrée central est une aube variable (118, 134).
  6. Moteur à turbine à gaz selon l'une quelconque des revendications 2 à 5, dans lequel le stator d'entrée central est chauffé pour antigivrage.
  7. Moteur à turbine à gaz selon l'une quelconque des revendications 2 à 6, dans lequel les une ou plusieurs pales d'inducteur (114, 132) sont chauffées pour antigivrage.
  8. Moteur à turbine à gaz selon une quelconque revendication précédente, dans lequel l'au moins un étage inducteur comprend une pluralité d'étages inducteurs (112, 130) accouplés au rotor de soufflante (106).
  9. Moteur à turbine à gaz selon la revendication 8, dans lequel chaque étage inducteur (112) comprend une ou plusieurs pales fixées pour rotation avec le rotor de soufflante (106) et un stator central fixé à une structure de moteur non rotative (36).
  10. Moteur à turbine à gaz selon la revendication 10, dans lequel le compresseur (44) est positionné juste à l'arrière de l'au moins un étage inducteur (112, 130).
  11. Moteur à turbine à gaz selon une quelconque revendication précédente, dans lequel l'au moins un étage inducteur (112) est positionné à l'avant du dispositif de réduction de vitesse (48).
  12. Moteur à turbine à gaz selon une quelconque revendication précédente, dans lequel le compresseur (44) comprend un compresseur basse-pression, et incluant un compresseur haute-pression (52) positionné à l'arrière du compresseur basse-pression (44) et entraîné par un second arbre (50) positionné radialement vers l'extérieur par rapport à l'arbre (40) qui entraîne le compresseur basse-pression (44).
  13. Moteur à turbine à gaz selon une quelconque revendication précédente, dans lequel le rotor de soufflante (106) tourne dans la même direction que l'arbre (40).
  14. Moteur à turbine à gaz selon l'une quelconque des revendications 1 à 12, dans lequel le rotor de soufflante (106) tourne dans une direction opposée à celle de l'arbre (40).
  15. Moteur à turbine à gaz selon une quelconque revendication précédente, le dispositif de changement de vitesse (48) comprend une boîte de vitesses incluant un planétaire (150) en prise par emboîtement avec des engrenages en étoile (152) et une couronne dentée (154) en prise par emboîtement avec les engrenages en étoile (152), et dans lequel la couronne dentée (154) entraîne la soufflante (42).
  16. Moteur à turbine à gaz selon l'une quelconque des revendications 1 à 14, dans lequel le dispositif de changement de vitesse (48) comprend une boîte de vitesses incluant un planétaire (160) en prise par emboîtement avec une pluralité de satellites (162) supportés par un porte-satellites (164) et une couronne dentée (166) en prise par emboîtement avec les satellites (162), et dans lequel la couronne dentée (166) est fixe et le porte-satellites (166) fournit une entrée dans la soufflante (42).
EP13784828.9A 2012-02-28 2013-02-13 Moteur à turbines à gaz équipé d'une section d'induction assemblée à la soufflante Active EP2820250B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/406,819 US9103227B2 (en) 2012-02-28 2012-02-28 Gas turbine engine with fan-tied inducer section
PCT/US2013/025802 WO2013165515A2 (fr) 2012-02-28 2013-02-13 Moteur à turbines à gaz équipé d'une section d'induction assemblée à la soufflante

Publications (3)

Publication Number Publication Date
EP2820250A2 EP2820250A2 (fr) 2015-01-07
EP2820250A4 EP2820250A4 (fr) 2015-12-09
EP2820250B1 true EP2820250B1 (fr) 2018-12-12

Family

ID=49003068

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13784828.9A Active EP2820250B1 (fr) 2012-02-28 2013-02-13 Moteur à turbines à gaz équipé d'une section d'induction assemblée à la soufflante

Country Status (4)

Country Link
US (1) US9103227B2 (fr)
EP (1) EP2820250B1 (fr)
SG (1) SG11201404761TA (fr)
WO (1) WO2013165515A2 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11560851B2 (en) 2012-10-02 2023-01-24 Raytheon Technologies Corporation Geared turbofan engine with high compressor exit temperature

Families Citing this family (45)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11149650B2 (en) 2007-08-01 2021-10-19 Raytheon Technologies Corporation Turbine section of high bypass turbofan
US8844265B2 (en) 2007-08-01 2014-09-30 United Technologies Corporation Turbine section of high bypass turbofan
US11242805B2 (en) 2007-08-01 2022-02-08 Raytheon Technologies Corporation Turbine section of high bypass turbofan
US11486311B2 (en) 2007-08-01 2022-11-01 Raytheon Technologies Corporation Turbine section of high bypass turbofan
US11346289B2 (en) 2007-08-01 2022-05-31 Raytheon Technologies Corporation Turbine section of high bypass turbofan
US20150377123A1 (en) 2007-08-01 2015-12-31 United Technologies Corporation Turbine section of high bypass turbofan
US9631558B2 (en) 2012-01-03 2017-04-25 United Technologies Corporation Geared architecture for high speed and small volume fan drive turbine
US9523422B2 (en) 2011-06-08 2016-12-20 United Technologies Corporation Flexible support structure for a geared architecture gas turbine engine
US9239012B2 (en) 2011-06-08 2016-01-19 United Technologies Corporation Flexible support structure for a geared architecture gas turbine engine
EP2610463B1 (fr) 2011-12-30 2016-08-03 United Technologies Corporation Train d'engrenages de moteur à turbine à gaz
US20150204238A1 (en) * 2012-01-31 2015-07-23 United Technologies Corporation Low noise turbine for geared turbofan engine
US9850821B2 (en) 2012-02-28 2017-12-26 United Technologies Corporation Gas turbine engine with fan-tied inducer section
US8915700B2 (en) 2012-02-29 2014-12-23 United Technologies Corporation Gas turbine engine with fan-tied inducer section and multiple low pressure turbine sections
US10125693B2 (en) 2012-04-02 2018-11-13 United Technologies Corporation Geared turbofan engine with power density range
EP2904254B8 (fr) * 2012-10-02 2020-11-04 Raytheon Technologies Corporation Turboréacteur double-flux à engrenages présentant une température de sortie de compresseur élevée
EP3090160A4 (fr) * 2013-12-30 2016-12-28 United Technologies Corp Moteur à turbine comprenant un nombre d'étages basse-pression équilibré
EP2899389A1 (fr) * 2014-01-27 2015-07-29 United Technologies Corporation Structure de support flexible pour moteur de turbine à gaz à architecture à engrenages
US10094277B2 (en) 2014-06-20 2018-10-09 United Technologies Corporation Gas turbine engine configured for modular assembly/disassembly and method for same
US10287976B2 (en) * 2014-07-15 2019-05-14 United Technologies Corporation Split gear system for a gas turbine engine
US20160047305A1 (en) * 2014-08-15 2016-02-18 General Electric Company Multi-stage axial compressor arrangement
EP3034833A1 (fr) * 2014-12-16 2016-06-22 United Technologies Corporation Moteur à turbine comprenant un comptage à étage basse pression équilibrée
US11067005B2 (en) 2015-02-03 2021-07-20 Raytheon Technologies Corporation Fan drive gear system
EP3112649B1 (fr) * 2015-07-01 2022-03-16 Raytheon Technologies Corporation Moteur à turbine à gaz avec section inducteur attachée à une soufflante
US10190599B2 (en) 2016-03-24 2019-01-29 United Technologies Corporation Drive shaft for remote variable vane actuation
US10443430B2 (en) 2016-03-24 2019-10-15 United Technologies Corporation Variable vane actuation with rotating ring and sliding links
US10294813B2 (en) 2016-03-24 2019-05-21 United Technologies Corporation Geared unison ring for variable vane actuation
US10443431B2 (en) 2016-03-24 2019-10-15 United Technologies Corporation Idler gear connection for multi-stage variable vane actuation
US10329946B2 (en) 2016-03-24 2019-06-25 United Technologies Corporation Sliding gear actuation for variable vanes
US10329947B2 (en) 2016-03-24 2019-06-25 United Technologies Corporation 35Geared unison ring for multi-stage variable vane actuation
US10107130B2 (en) 2016-03-24 2018-10-23 United Technologies Corporation Concentric shafts for remote independent variable vane actuation
US10458271B2 (en) 2016-03-24 2019-10-29 United Technologies Corporation Cable drive system for variable vane operation
US10415596B2 (en) 2016-03-24 2019-09-17 United Technologies Corporation Electric actuation for variable vanes
US10288087B2 (en) 2016-03-24 2019-05-14 United Technologies Corporation Off-axis electric actuation for variable vanes
US10301962B2 (en) 2016-03-24 2019-05-28 United Technologies Corporation Harmonic drive for shaft driving multiple stages of vanes via gears
US10544734B2 (en) 2017-01-23 2020-01-28 General Electric Company Three spool gas turbine engine with interdigitated turbine section
US10539020B2 (en) 2017-01-23 2020-01-21 General Electric Company Two spool gas turbine engine with interdigitated turbine section
US10655537B2 (en) 2017-01-23 2020-05-19 General Electric Company Interdigitated counter rotating turbine system and method of operation
US10544793B2 (en) 2017-01-25 2020-01-28 General Electric Company Thermal isolation structure for rotating turbine frame
US10876407B2 (en) 2017-02-16 2020-12-29 General Electric Company Thermal structure for outer diameter mounted turbine blades
US10294821B2 (en) 2017-04-12 2019-05-21 General Electric Company Interturbine frame for gas turbine engine
US10787931B2 (en) 2017-05-25 2020-09-29 General Electric Company Method and structure of interdigitated turbine engine thermal management
US10669893B2 (en) 2017-05-25 2020-06-02 General Electric Company Air bearing and thermal management nozzle arrangement for interdigitated turbine engine
US10605168B2 (en) 2017-05-25 2020-03-31 General Electric Company Interdigitated turbine engine air bearing cooling structure and method of thermal management
US10718265B2 (en) 2017-05-25 2020-07-21 General Electric Company Interdigitated turbine engine air bearing and method of operation
US11428160B2 (en) 2020-12-31 2022-08-30 General Electric Company Gas turbine engine with interdigitated turbine and gear assembly

Family Cites Families (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3903690A (en) * 1973-02-12 1975-09-09 Gen Electric Turbofan engine lubrication means
GB2259328B (en) 1991-09-03 1995-07-19 Gen Electric Gas turbine engine variable bleed pivotal flow splitter
US6647708B2 (en) * 2002-03-05 2003-11-18 Williams International Co., L.L.C. Multi-spool by-pass turbofan engine
US7334392B2 (en) * 2004-10-29 2008-02-26 General Electric Company Counter-rotating gas turbine engine and method of assembling same
EP1834071B1 (fr) 2004-12-01 2013-03-13 United Technologies Corporation Inducteur de pale de ventilateur de moteur de turbine a pression d'entree
EP1841959B1 (fr) 2004-12-01 2012-05-09 United Technologies Corporation Ailettes de rotor de soufflante pour moteur a turbine en bout
US7927075B2 (en) 2004-12-01 2011-04-19 United Technologies Corporation Fan-turbine rotor assembly for a tip turbine engine
US7726113B2 (en) * 2005-10-19 2010-06-01 General Electric Company Gas turbine engine assembly and methods of assembling same
US7694505B2 (en) 2006-07-31 2010-04-13 General Electric Company Gas turbine engine assembly and method of assembling same
US7966806B2 (en) * 2006-10-31 2011-06-28 General Electric Company Turbofan engine assembly and method of assembling same
US7716914B2 (en) 2006-12-21 2010-05-18 General Electric Company Turbofan engine assembly and method of assembling same
US9957918B2 (en) 2007-08-28 2018-05-01 United Technologies Corporation Gas turbine engine front architecture
US8015798B2 (en) 2007-12-13 2011-09-13 United Technologies Corporation Geared counter-rotating gas turbofan engine
US8887485B2 (en) 2008-10-20 2014-11-18 Rolls-Royce North American Technologies, Inc. Three spool gas turbine engine having a clutch and compressor bypass
US8166748B2 (en) * 2008-11-21 2012-05-01 General Electric Company Gas turbine engine booster having rotatable radially inwardly extending blades and non-rotatable vanes
US20110167791A1 (en) 2009-09-25 2011-07-14 James Edward Johnson Convertible fan engine
US8672801B2 (en) 2009-11-30 2014-03-18 United Technologies Corporation Mounting system for a planetary gear train in a gas turbine engine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None *

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11560851B2 (en) 2012-10-02 2023-01-24 Raytheon Technologies Corporation Geared turbofan engine with high compressor exit temperature

Also Published As

Publication number Publication date
US20130224003A1 (en) 2013-08-29
EP2820250A2 (fr) 2015-01-07
US9103227B2 (en) 2015-08-11
WO2013165515A3 (fr) 2014-01-23
EP2820250A4 (fr) 2015-12-09
WO2013165515A2 (fr) 2013-11-07
SG11201404761TA (en) 2014-09-26

Similar Documents

Publication Publication Date Title
EP2820250B1 (fr) Moteur à turbines à gaz équipé d'une section d'induction assemblée à la soufflante
EP2820249B1 (fr) Moteur à turbine à gaz avec section aubage d'alimentation associée à une soufflante et plusieurs sections de turbine basse pression
EP2877725B1 (fr) Soufflante à réducteur dotée d'un compresseur contrarotatif interne
US11143111B2 (en) Fan drive gear system mechanical controller
US9816442B2 (en) Gas turbine engine with high speed low pressure turbine section
EP2820280B1 (fr) Turbine basse pression à rotation inverse avec système d'engrenage monté sur le carter d'échappement de turbine
EP3101258B1 (fr) Architecture à engrenages pour un moteur à turbine à gaz et procédé correspondant
EP3036416B1 (fr) Moteur à turbine à gaz à engrenages de poussée élevée
US9194290B2 (en) Counter-rotating low pressure turbine without turbine exhaust case
US9850821B2 (en) Gas turbine engine with fan-tied inducer section
EP2932073B1 (fr) Turbocompresseur pour air de prélèvement
EP2809912A1 (fr) Moteur à turbosoufflante à engrenages doté d'arbre contrarotatifs
EP3054141B1 (fr) Réduction d'engrenage pour turboréacteur à engrenage
EP2820266A2 (fr) Turbine basse pression contrarotative à système d'engrenage monté sur un carter de sortie turbine
US20160061052A1 (en) Gas turbine engine with high speed low pressure turbine section
EP3112649B1 (fr) Moteur à turbine à gaz avec section inducteur attachée à une soufflante
EP2809937B1 (fr) Agencement de palier d'arbre de moteur de turbine à gaz
EP3081768B1 (fr) Configuration de palier d'arbre de moteur à turbine à gaz
EP2955325B1 (fr) Turboréacteur à engrenages à rotor à aubage intégral

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 20140925

AK Designated contracting states

Kind code of ref document: A2

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

AX Request for extension of the european patent

Extension state: BA ME

DAX Request for extension of the european patent (deleted)
A4 Supplementary search report drawn up and despatched

Effective date: 20151110

RIC1 Information provided on ipc code assigned before grant

Ipc: F02C 7/047 20060101ALI20151104BHEP

Ipc: F01D 5/02 20060101ALI20151104BHEP

Ipc: F01D 15/12 20060101AFI20151104BHEP

Ipc: F02K 3/072 20060101ALI20151104BHEP

Ipc: F02C 7/36 20060101ALI20151104BHEP

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: UNITED TECHNOLOGIES CORPORATION

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: GRANT OF PATENT IS INTENDED

INTG Intention to grant announced

Effective date: 20180620

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE PATENT HAS BEEN GRANTED

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

REG Reference to a national code

Ref country code: GB

Ref legal event code: FG4D

REG Reference to a national code

Ref country code: CH

Ref legal event code: EP

REG Reference to a national code

Ref country code: AT

Ref legal event code: REF

Ref document number: 1076262

Country of ref document: AT

Kind code of ref document: T

Effective date: 20181215

REG Reference to a national code

Ref country code: DE

Ref legal event code: R096

Ref document number: 602013048165

Country of ref document: DE

REG Reference to a national code

Ref country code: IE

Ref legal event code: FG4D

REG Reference to a national code

Ref country code: NL

Ref legal event code: MP

Effective date: 20181212

REG Reference to a national code

Ref country code: LT

Ref legal event code: MG4D

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: BG

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190312

Ref country code: ES

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: HR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: LT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: LV

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: NO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190312

Ref country code: FI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

REG Reference to a national code

Ref country code: AT

Ref legal event code: MK05

Ref document number: 1076262

Country of ref document: AT

Kind code of ref document: T

Effective date: 20181212

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: AL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: RS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: GR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190313

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: NL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CZ

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: PT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190412

Ref country code: IT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: PL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: RO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: IS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190412

Ref country code: SM

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: EE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 602013048165

Country of ref document: DE

REG Reference to a national code

Ref country code: CH

Ref legal event code: PL

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: LU

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190213

Ref country code: AT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: SI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: MC

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

Ref country code: DK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

26N No opposition filed

Effective date: 20190913

REG Reference to a national code

Ref country code: BE

Ref legal event code: MM

Effective date: 20190228

REG Reference to a national code

Ref country code: IE

Ref legal event code: MM4A

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CH

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190228

Ref country code: LI

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190228

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190213

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: BE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190228

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: TR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190213

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CY

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: HU

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT; INVALID AB INITIO

Effective date: 20130213

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181212

REG Reference to a national code

Ref country code: DE

Ref legal event code: R081

Ref document number: 602013048165

Country of ref document: DE

Owner name: RAYTHEON TECHNOLOGIES CORPORATION (N.D.GES.D.S, US

Free format text: FORMER OWNER: UNITED TECHNOLOGIES CORPORATION, FARMINGTON, CONN., US

P01 Opt-out of the competence of the unified patent court (upc) registered

Effective date: 20230520

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: DE

Payment date: 20240123

Year of fee payment: 12

Ref country code: GB

Payment date: 20240123

Year of fee payment: 12

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 20240123

Year of fee payment: 12